In conventional gas turbine combustors liquid or gaseous fuel is sprayed directly into the combustion chamber for combustion in admixture with air. Consequently, fuel-air mixing and vaporization occur in the combustion zone resulting in significant regions of stoichiometric combustion and high NO.sub.x formation.
Accordingly, to achieve lower NO.sub.x levels there has been interest in and much development of premixed/prevaporized fuel combustion systems such the dry low NO.sub.x natural gas combustors now used commercially. However, such combustors not only can have stability problems stemming from the need to operate near the lean limit but as with any premixed/prevaporized combustion system there is the potential for propagation of the flame upstream to the mixing/vaporization region with resultant stoichiometric burning and damage to the combustor. Although a safety shut-off can minimize damage, a shut down and inspection would be essential. With liquid fuels the problem is even greater. Moreover, the high combustor inlet temperatures not only of advanced stationary and aero gas turbine designs but even of most present day aero engines greatly increase the likelihood of such an occurrence. The problem is so severe that it has been questionable as to whether any premixed combustor will ever be feasible for an aero engine inasmuch as no conventional device has been deemed adequate to avoid engine damage. Not only must a device be able to block upstream flame propagation but it must impose a negligible pressure drop, i.e., less than about one percent. The present invention offers a practical low pressure drop solution to this important problem.